The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzles therein.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. A high pressure turbine (HPT) follows the combustor for extracting energy from the combustion gases to power the compressor. A low pressure turbine (LPT) follows the HPT and extracts additional energy from the gases to power an upstream fan in the typical turbofan gas turbine engine application. Alternatively, the LPT may drive an external drive shaft for marine and industrial applications.
The HPT includes one or more turbine nozzles for directing the combustion gases into corresponding stages of turbine rotor blades extending radially outwardly from supporting rotor disks. The nozzle stator vanes and the turbine rotor blades are typically hollow and contain internal cooling circuits therein through which is circulated pressurized air bled from the compressor for use in cooling the metal material thereof during operation.
The art is crowded with various configurations for cooling the nozzle stator vanes and turbine rotor blades and their adjacent components which are subject to heating by the hot combustion gases which flow through the turbine flowpaths during operation. For example, the cooling air is bled from the compressor and channeled through one circuit along the rotor disks and into the individual rotor blades through inlets provided in the mounting dovetails thereof supported in the perimeter of the rotor disks.
The individual blade airfoils typically have multiple radial flow passages therein for providing internal impingement cooling or serpentine cooling, or both, inside the airfoil typically with small turbulator ribs disposed along the inner surface of the airfoil for tripping the cooling flow to increase its heat transfer performance. The spent cooling air is discharged through various rows of film cooling holes extending through the pressure and suction sides, or both, of the airfoil in various configurations.
Correspondingly, additional air may also be bled from the compressor in another circuit to supply the turbine nozzles with cooling air, typically through their outer bands. The first stage turbine nozzle is subjected to the hottest temperature combustion gases from the combustor and typically includes multiple cooling passages or cavities inside the individual nozzle vanes with associated internal impingement baffles for increasing the cooling performance of the pressurized air.
In a two stage HPT, the second stage nozzle may also be internally cooled by the compressor bleed air, and may also include an internal impingement baffle therein for enhancing cooling performance.
In both nozzle stages, the spent cooling air is typically discharged through various rows of film cooling holes in the pressure or suction sides, or both, of the vane airfoils for return to the main combustion gas flowpath during operation.
Since the second stage turbine nozzle is disposed axially between the first and second stage rotor blades and their corresponding rotor disks, a portion of the cooling air channeled through the vanes is typically discharged through the inner band of the nozzle for providing purge cooling flow in the forward and aft cavities defined with the corresponding rotor disks of the first and second stages. A honeycomb seal is typically supported from the inner band and cooperates with an annular seal having labyrinth seal teeth disposed closely adjacent thereto, with the forward and aft rotor cavities being defined on the opposite sides of the labyrinth seal bounded by the corresponding rotor disks.
In this way, cooling air channeled through the nozzle vanes may be discharged into the forward rotor cavity for cooling the aft face of the first stage rotor with the spent purge air leaking past the labyrinth seal teeth for then purging the aft rotor cavity and correspondingly cooling the forward face of the second stage rotor disk.
The various cooling configurations provided for the stator vanes, turbine blades, purge cavities, and other components bounding the hot combustion gas flowpath are typically tailored to the different operating environments thereof for maximizing cooling effectiveness while minimizing the use of compressor bleed air. Any air bled from the compressor which is not used in the combustion process decreases the overall efficiency of the gas turbine engine and requires more fuel burn.
In aircraft engine applications fuel consumption is always a paramount design objective, with modern aircraft turbofan engines designed therefor being constantly improved for minimizing fuel consumption. Since the HPT is subject to the hottest combustion gases during operation, reducing cooling air requirements therefor is a significant challenge in view of the hostile temperature environment of this section of the engine, and in view of the mature design thereof.
Accordingly, it is desired to further improve cooling efficiency in high pressure turbines for correspondingly further improving performance and efficiency of the gas turbine engine.